Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. The turbine blade assembly usually comprises a number of rings of turbine blades between which nozzle arrangements comprising a number of guide vanes are located.
A nozzle arrangement typically comprises an outer carrier ring or support ring, an inner carrier ring or support ring, and a number of nozzle segments each typically comprising a radial outer platform, a radial inner platform and at least one vane extending from the radial outer platform to the radial inner platform. The nozzle arrangement forms an annular flow path for hot and corrosive combustion gases from the combustor.
Combustors often operate at high temperatures that may exceed 1350° C. Typical turbine combustor configurations expose turbine vane and blade arrangements to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for prolonging the lifetime of the vanes and the blades and for reducing the likelihood of failure as a result of excessive temperatures.
In order to prevent the platforms of the nozzle segments, which form the walls of the flow path for the hot and corrosive combustion gases, from damage due to the hot combustion gases the platforms are cooled with compressor air. However, the pressure of the compressor air used for cooling the platforms is higher than the pressure of the combustion gases flowing downstream of the nozzle arrangement. Moreover, the cooling air used for cooling the platforms, in particular their downstream ends, will be discharged into the flow path of the hot combustion gases. Hence, the flow of air into the flow path needs to be restricted to a minimum in order to preserve overall turbine efficiency. In order to restrict the flow of compressor air into the flow path of the hot combustion gas seals are provided between the radial outer platform of the nozzle segments and the outer carrier ring. Moreover, seals are provided between the radial inner platform of the nozzle segments and the inner carrier ring, mainly for preventing hot combustion gas from entering gaps between the platform and the carrier ring. Examples of such seals are disclosed in US 2008/0101927 A1, U.S. Pat. Nos. 6,641,144, 6,572,331, 6,637,753, 6,637,751, US 2005/0244267 A1, EP 1 323 890 B1, EP 1 323 896 B1, EP 1 323 898 B1, U.S. Pat. No. 6,752,331, and US 2003/012398 A1.
EP 1 247 942 B1 is further disclosing a seal element for sealing a gas-path leakage-gap between components of a turbo machinery. This seal element consists of a plurality of elements made of sheet metal with of ceramic material. US 2005/0095123 A1does disclose a segmented seal between two longitudinally adjacent elements of a turbo machine.
U.S. Pat. No. 4,126,405 discloses a turbine nozzle with a leaf seal located between a vane forward outer rail and a combustor rear flange. The leaf seal is held in place by a plurality of pins by which it is fixed to the outer rail of the vane.
WO 00/77348 A1 describes a gas turbine with a reverse airflow duct between a combustion chamber and a first nozzle stage of the turbine. An inner duct wall of the reverse airflow duct is an integrally cast extension of a nozzle shroud and is covered by an impingement blade which allows for impingement cooling of the duct wall. A sealing lip is present between the duct wall and the inner combustor wall.
Known sealing devices do need a complex fastening means to mount them on a nozzle arrangement. All known sealing arrangements further do have in common that their construction, assembly, and manufacturing costs due to complexity are relatively high.